This invention relates to improvements in turbine nozzles used in turbine engines having centrifugal compressors. Of particular concern is the elimination of combustor hot spot effects on turbine nozzles associated with small and medium sized engines.
It is known to make diffusers which include a rotary stage. U.S. Pat. No. 3,941,501 to Shank discloses a first stage vaneless diffuser having rotating sidewalls which freely turn on bearings mounted coaxially with the compressor. U.S. Pat. No. 3,868,196 to Lown shows a rotating vaneless diffuser which is powered by leakage flow from the impeller disk of the compressor. Both of the above patents disclose means for efficiently matching a diffuser to a compressor rotor which delivers gas at supersonic velocity. They do this by introduction of a rotating stage having sidewalls which travel at about half that of the impeller. In this way, gas molecules impact the diffuser sidewalls at subsonic velocities thereby minimizing shock wave phenomena in the diffuser.
My rotary diffuser serves a different purpose. Torque is applied to my diffuser stage in an amount adequate to rotate a first nozzle disk positioned at the outlet of the engine combustors. By spinning the turbine nozzle system disk at a nominal rate, hot spot effects do not develop on those nozzle blades which are directly in front of the combustor exit. Rather, each balde would spend about 20 milli-seconds in the hottest part of the flame before moving into a slightly cooler environment. This reduces the total heat transfer rate to the vane. Consequently, the hot spot effect is eliminated or minimized.